Disk performance by co-extrusion

ABSTRACT

A method for fabricating a gas turbine engine disk by co-extrusion is described which comprises the steps of preparing a cylindrical shaped preform billet having a central bore region of preselected radius r and comprising a first material, and a rim region surrounding the bore region to a radius R of the preform billet and comprising a second material, extruding the preform billet at preselected extrusion temperature through an extrusion die to form an extrusion product of preselected reduction, and removing the disk from the extrusion product.

RIGHTS OF THE GOVERNMENT

The invention described herein may be manufactured and used by or forthe Government of the United States for all governmental purposeswithout the payment of any royalty.

BACKGROUND OF THE INVENTION

The present invention relates generally to methods for the fabricationof high temperature strength engine components, and more particularly toa co-extrusion method for manufacturing gas turbine engine disks forhigh temperature operation.

In the operation of gas turbine engines, temperatures encountered byengine components require that a rotating turbine disk within the enginepreferably exhibit high temperature resistance to creep and stressrupture in the rim of the disk, and high temperature ultimate tensilestrength in the bore. Typical rotating engine components which areexposed to high operating temperatures and wherein strengthcharacteristics in the rim different from that of the bore are desirableinclude components in all stages of the compressor spool and all otherturbine disks exposed to hot gas flow downstream of the combustor regionof the engine. Prior art fabrication methods for engine disks havingdifferent rim/bore materials consist chiefly of inertia welding a boreor web of preselected material, size and configuration to a ring ofdifferent material and corresponding size comprising the rim of thefinished disk. In the inertia welding process, the bore/web is rotatedat preselected high speed and pushed against the stationary rim togenerate by friction heat required to make a weld between bore and rim.The heat affected zone in such welds is characteristically narrow ascompared to that obtained by fusion welding processes. Engine disksfabricated by the inertia welding process chiefly consist of bores oftitanium base and nickel base superalloy fatigue (low and high) and/orburst (high ultimate tensile strength) resistant type metals and alloys,including Ti6-4, Ti-17, Ti6-2-4-6, Rene-95, Inco-718, Merl-76, In-100,and Rene-88DT, and rims of respective similar metals and alloys,including Ti6-2-4-2, IMI-829, Astroloy, R-88DT, In-100, U-500, U-700,U-720, and Rene-41, wherein a respective pair of materials must beselected for amenability to welding by the inertia process. The inertiawelding process for fabricating engine disks therefore suffers fromcertain shortcomings including limitation on bore/rim material selectionto those combinations amenable to inertia welding. undesirably lowattainable heat affected zone/weld thickness between bore and rim. lowstrength or embrittlement of the material in the bore/rim weldinterface, and machine oapacity limitations due to large weld surfaceareas and/or large diameters.

The invention solves or reduces in critical importance problems withprior methods associated with the fabrcation of engine disks for turbineengines by providing a co-extrusion method for fabricating gas turbineengine disks having high temperature resistance to creep and stressrupture in the rim and high tensile strength and/or high fatigue (low)or crack growth rate resistance in the bore or web, wherein the twomaterials comprising. respectively, the rim and bore of the disk aresimultaneously co-extruded at a appropriate preselected extrusiontemperature from a preform billet of the two materials, resulting in asolid state metallurgical bond between the two materials, the extrusionbeing subsequently processed thermomechanically (e.g. by forging) and/ormachined to achieve the desired shape or to perform appropriatemechanical work reductions to desired product size. The method of theinvention allows formation of disks from substantially any combinationof materials desired for the rim and bore, respectively. and,accordingly, preferred combinations of titanium base and nickel basesuperalloy fatigue (low and high) and/or burst (high ultimate tensilestrength) resistant type metals and alloys, including Ti6-4, Ti-17,Ti6-2-4-6, Rene-95, Rene-88DT, In-400, and Merl-76, for the bore andcreep rupture resistance type metals and alloys, including Ti6-2-4-2,IMI-829, Udimet-500, Udimet-700, Udimet-720, and Astroloy, for the rimmay be used, many of which which were not practical using prior diskfabrication methods (e.g., inertia welding). The preform billet maycomprise any combination of wrought, powder or cast material in anycombination of alloys in any number of concentric layers (rings) toachieve desired product characteristics. Components fabricated fromextrusions made according to the invention have improved and predictablemetallurgical and structural properties as compared to those made by themore expensive inertia welding process.

It is therefore a principal object of the invention to provide animproved method for fabricating high temperature resistant gas turbineengine disks.

It is a further object of the invention to provide a method forfabricating gas turbine disks by co-extrusion.

It is a further object of the invention to provide an improved gasturbine engine disk fabricated by co-extrusion.

These and other objects of the invention will become apparent a thedetailed description of representative embodiments proceeds.

SUMMARY OF THE INVENTION

In accordance with the foregoing principles and objects of theinvention, a method for fabricating a gas turbine engine disk byco-extrusion is described which comprises the steps of preparing acylindrical shaped preform billet having a central bore region ofpreselected radius r and comprising a first material, and a rim regionsurrounding the bore region to a radius R of the preform billet andcomprising a second material, extruding the preform billet atpreselected extrusion temperature through an extrusion die to form anextrusion product of preselected reduction, and removing the disk fromthe extrusion product.

DESCRIPTION OF THE DRAWINGS

The invention will be clearly understood from the following detaileddescription of representative embodiments thereof read in conjunctionwith the accompanying drawings wherein:

FIG. 1 is a schematic perspective view, partially in section, of anextrusion die and co-extruded preform billet of the invention;

FIG. 2 is a view along line A--A of the preform billet of FIG. 1; and

FIG. 3 is a view along line B--B of the extruded billet of FIG. 1.

DETAILED DESCRIPTION

Referring now to the drawings, shown in FIG. 1 is a schematicperspective view, partially in section, of a co-extrusion preform billet10 extruded under the influence of a conventionally applied force 11through extrusion die 13 according to the method of the invention.Referring additionally to FIG. 2, which is a sectional view along lineA--A of preform billet 10, it is seen that preform billet 10 comprisestwo or more components 15, 17 in the form of a cylindrical borecomponent 15 and rim component 17. As discussed above, bore component 15is selected to provide high ultimate tensile strength in the centralregion of the finished engine disk and, accordingly, may preferablycomprise Ti6-4, Ti-17, Ti6-2-4-6, Rene-95, Rene-88DT, Inco-718, In-100,or Merl-76, or similar material as would occur to one with skill in thefield of the invention. Rim component 17 is selected to proVide hightemperature resistance to creep and stress rupture in the rim region ofthe finished engine disk and, accordingly, may preferably comprise Ti6-2-4-2, IMI-829, Astroloy, Udimet-500, Udimet-700, Udimet-720, Rene-41, orMerl-76. Components 15, 17 may be in any suitable form such as wroughtmetal, RST powder, casting or other form with substantially any heattreat history, the same not limiting the invention herein. Further,additional rings of metals/alloys in selected thicknesses may beincluded between components 15, 17 for attaining any desirable radialvariation in structural or metallurgical properties in the finishedproduct engine disk. All components 15,17 may be contained in anextrusion can 19 of appropriate size and material. For most extrusionsperformed in demonstration of the invention. extrusion cans of aluminum,copper, stainless steel, or carbon steel having suitable wall thicknessin the range of about 0.25 to about 1.0 inch proved satisfactory.

Preform 10 may be preheated conventionally as by heating means 21 to atemperature appropriate for the intended extrusion through die 13. Die13 may also be heated conventionally as by heating means 23 to providedesirable preselected reduction of diameter of preform billet 10 to thatof extrusion 25. The temperature at which preform billet 10 is extrudeddepends on material selections for components 15,17. Co-extrusionsperformed in demonstration of the invention showed that in selecting anappropriate extrusion temperature for preform billet 10, considerationof the melting point of the material comprising the bore component 15was most critical in obtaining extrusions 25 of the desired rimthickness and overall outside diameter. Accordingly, optimum extrusiontemperatures were in the range of from about 50 percent to about 90percent, and preferably about 85 percent, of the melting point in °K. ofthe bore component 15.

Referring now additionally to FIG. 3, shown therein is a sectional viewalong line B--B of extrusion 25, which, taken with the sectional view ofFIG. 2. illustrates the reduction of bore component 15 radius r inpreform billet 10 to r' in component 15' of extrusion 25, of rimcomponent 23 thickness (R-r) in preform billet 10 to (R'-r') incomponent 17' of extrusion 25, and of the reduction in thickness ofextrusion can 19 as a result of the extrusion through die 13. The radiusof bore component 15 and the thicknesses of rim component 17 and of anyadditional rings of material therebetween may be selected according todesired end product specifications. Engine disks fabricated of thepreferred components 15, 17 listed above have a ratio R'/r' in the rangeof from about 1.1 to about 2.5. Accordingly, in extrusions made indemonstration of the invention, best results were obtained wherein theratio of R/r in preform billet 10 was from about 2.0 to about 2.5, andextrusions were made at the selected temperature defined above throughdie 13 having an extrusion ratio of about 3:1 to about 8:1.

Extrusion 25 may be subjected to more than one stepwise extrusionprocess to provide a desired overall reduction from the initial preformbillet 10 diameter R to the finished product having the desired radialvariation of component materials. Once extrusion 25 of the desiredreduction (diameter R') is obtained, extrusion can 19' is conventionallystripped therefrom and the product comprising the extruded bore and rimcomponents 15',17' is worked by forging, heat treating, or otherwise toend product specifications and cut as at 27 to define a finished disk29. Disk 29 is then further treated and/or machined to intended size andperipheral shape of the completed turbine engine disk or like component.

The invention therefore provides a novel inexpensive method forfabrication of high temperature operating gas turbine engine disks byco-extrusion. Engine disks having superior structural and metallurgicalproperties may be fabricated from two separate metals or alloys whereina solid state metallurgical bond is achieved therebetween withoutwelding. It is understood that modifications to the invention may bemade as might occur to one with skill in the field of the inventionwithin the scope of the claims. All embodiments contemplated hereunderwhich accomplish the objects of the invention have therefore not beenshown in complete detail. Other embodiments may be developed withoutdeparting from the spirit of the invention or from the scope of theclaims.

We claim:
 1. A method for fabricating a high temperature resistant diskfor a gas turbine engine, comprising the steps of:(a) preparing acylindrical shaped preform billet from at least two materials, saidpreform billet having a central bore region defined along a central axisof said preform billet, said bore region being of preselected radius rand comprising a first material of said at least two materials and a rimregion surrounding said bore region to a preselected radius R of saidpreform billet, said rim region comprising a second material of said atleast two materials; (b) extruding said preform billets at preselectedextrusion temperature through an extrusion die to form an extrusionproduct of preselected reduction; and (c) cutting said extrusion productto preselected axial thickness to define a disk of preselected diameterand thickness.
 2. The method of claim 1 wherein said extrusion isperformed at a temperature in the range of from about 50 to 90 percentof the melting point in °K. of said first material.
 3. The method ofclaim 2 wherein said extrusion is performed at a temperaturecorresponding to about 85 percent of the melting point in °K. of saidfirst material.
 4. The method of claim 1 wherein the extrusion ratio ofsaid die is between 3:1 and 8:1.
 5. The method of claim 2 wherein theratio R:r is selected in the range of from 1.1 to 2.5.
 6. The method ofclaim 1 wherein said first material is selected from the groupconsisting of Ti6-4, Ti-17, Ti6-2-4-6, Rene-95, Inco-718, Rene-88DT, andIn-100, and the second material is selected from the group consisting ofTi6-2-4-2, IMI-829, Astroloy, U-500, U-700, U-720, Rene-41, and Merl-76.7. The method of claim 1 wherein said preform billet is contained in anextrusion can.
 8. The method of claim 7 wherein said extrusion cancomprises a material selected from the group consisting of aluminum,copper, stainless steel, and carbon steel.
 9. The method of claim 8wherein said extrusion can has wall thickness of 0.25 to 1.0 inch.